As is well known in the gas turbine engine art, it is manifestly important to maximize the use of compressor air that is utilized outside the engine cycle. Of particular importance is the use of compressor air utilized to cool the turbine blades and to assure that the lower pressurized air is used rather than air that is at a higher pressure. Obviously, the lower the pressure of the air being used for turbine blade cooling, the lower the performance penalty and the overall improvement in engine performance. Additionally, utilizing a lower pressure improves the designer's ability to reduce leakages. And the lower pressure air is cooler and hence more effective for cooling purposes.
One aspect that contributes to the higher pressure of the compressor air is the fact that a predetermined pressure ratio across the turbine film cooled holes is necessary to obtain adequate film cooling of the exit air. By utilizing the total pressure instead of the static pressure of the cooling air for feeding the impingement cavities and increasing the outflow margin, i.e., the pressure ratio across the stagnation point row of film holes, will permit the use of a lower supply pressure (compressor air).